This invention relates to gas turbines and, more particularly, to a concept for attaching tubes associated with cooling systems for ultrahigh temperature turbine rotor blades to the rotating turbine disk.
It is well understood that gas turbine engine shaft horsepower and specific fuel consumption, which is the rate of fuel consumption per unit of power output, can be improved by increasing turbine inlet temperatures. However, current turbines are limited in inlet temperature by the physical properties of their materials. To permit turbines to operate at gas stream temperatures which are higher than the materials can normally tolerate, considerable effort has been devoted to the development of sophisticated methods of turbine cooling. In early turbine designs, cooling of high temperature components was limited to transferring heat to lower temperature parts by conduction, and air cooling was limited to passing relatively cool air across the face of the turbine rotor disks.
In order to take advantage of the potential performance improvements associated with even higher turbine inlet temperatures, modern turbine cooling technology utilizes hollow turbine nozzle vanes and blades to permit operation at inlet gas temperatures in the 2,000.degree. to 2,300.degree. F. (1,094.degree. to 1,260.degree. C.) range. Various techniques have been devised to air cool these hollow blades and vanes. These incorporate three basic forms of air cooling, either singly or in combination, depending on the level of gas temperatures encountered and the degree of sophistication permissible. These basic forms of air cooling are known as convection, impingement and film cooling. U.S. Pat. Nos. 3,700,348 and 3,715,170, assigned to the assignee of the present invention, are excellent examples of advanced turbine air-cooling technology incorporating these basic air-cooling forms.
However, the benefits obtained from sophisticated air-cooling techniques are at least partially offset by the extraction of the necessary cooling air from the propulsive cycle. For example, probably the most popular turbine coolant today is air which is bled off of the compressor portion of the gas turbine engine and is routed to the hollow interior of the turbine blades. Typically, the work which has been done on this air by the compressor is partially lost to the cycle. Additionally, as the cooling air circulates throughout the turbine blade it picks up heat from the metallic blades or vanes. When this heat cooling air leaves the turbine blades, perhaps as a coolant film, this heat energy is lost to the cycle since the hot gases are normally mixed with the exhaust gases and ejected from an engine nozzle. It would be desirable, therefore, to have a cooling system wherein a medium other than compressor bleed air is used and wherein the heat extracted by the cooling medium is put back into the cycle in a useful and practical manner.
A partial solution to the foregoing problems has been the suggestion of closed-loop cooling systems for turbine blades which may or may not also incorporate the concept of regeneration or recuperation to recover lost thermal energy. However, such closed-loop systems typically comprise a fluid coolant reservoir at or near the turbine disk or shaft rotational axis and conduits are provided to route the coolant to and through the blades. In some systems, the coolant conduits comprise radial bores through the disk. Such an arrangement is taught in U.S. Pat. Nos. 2,778,601 and 3,756,020. The disadvantage of such devices is that the disk passages tend to degrade the disk structural integrity, an important consideration in aircraft gas turbine engine design, and tend to increase its cost.
U.S. Pat. No. 2,849,210 teaches a turbine using the closed-loop thermosiphon principle wherein the hollow blade interiors are fluidly connected to an annular condensing chamber near the disk rotational center by a plurality of tubes extending down one side of the disk between a manifolded vaporizing chamber associated with the blade and the condensing chamber. However, the tubes are not attached to the turbine disk. It will be appreciated that in order to enhance structural integrity of the entire rotating system and to prevent stresses from arising in the tubes which would tend to unnecessarily bend, twist or stretch them, it is more desirable to provide a means for mechanically attaching the tubes to the rotating disk. Preferably, the method of mechanical attachment would permit easy assembly and replacement of the tubes and blades.
Accordingly, it is the primary object of the present invention to provide apparatus for attaching generally radially extending thin-walled tubes to the sides of a rotating turbine disk in order to preclude unnecessary stresses due to forces generated by high rotational speeds.
It is a further object of the present invention to provide a mechanical attachment for such tubes which will facilitate their assembly and replacement.
These, and other objects and advantages, will be more clearly understood from the following detailed descriptions, the drawings and specific examples, all of which are intended to be typical of rather than in any way limiting to the present invention.
Briefly stated, the above objectives are accomplished by providing a rotating member, such as a turbine disk, with an axially extending retaining means such as a ring having a plurality of D-shaped, open-ended slots formed about an edge thereof. Each tube is provided with a wear collar shaped to the contour of the slot into which it is received, and each collar, in turn, is provided with a pair of laterally extending lips which receive the ring therebetween and which limit radial movement of the collar. A circumferentially extending groove about the inner circumference of the ring intersects each groove and receives a locking means such as a split locking ring which limits axial movement of the collars and tubes.